Control moment gyros (CMG's) are widely used in attitude control systems of satellites and other space vehicles. The words “spacecraft”, “satellite” and “space vehicle” are used synonymously herein. Persons of skill in the art understand that multiple CMG's are often used to permit orientation control in three dimensions. However, for convenience of explanation, it is assumed herein that the orientation system is one dimensional, i.e., employing only one CMG. Persons of skill in the art will understand that multiple CMG's can be employed to provide for orientation in three dimensions and that the present invention applies to such multidimensional arrangements.
FIG. 1 is a simplified schematic block diagram illustrating conventional prior art spacecraft attitude control system 20 employing CMG 22 used to orient spacecraft 24 in space. Spacecraft Control System 26 decides what orientation that the satellite should assume and issues an appropriate command over link 261 to Attitude Control Processor (ACP) 28. ACP 28 takes into account information received from rate sensors 241 and attitude sensors 242 on spacecraft 24 concerning the current rate at which spacecraft 24 is rotating (if at all) and the orientation (attitude) of the axis of rotation, and then issues a rate command (e.g., in radians per second or other convenient rotational units) over link 29 to CMG 22 to cause spacecraft 24 to rotate to the new orientation desired by SCS 26. CMG 22 comprises a rotating mass held in a moveable gimbal which is in turn mechanically coupled to spacecraft 24. When the gimbal is rotated in response to the rate command received by CMG 22, the gimbal processes thereby imparting torque 25 to spacecraft 24 causing it to begin rotating since the total momentum of the overall system is conserved. Based on the feedback to ACP 28 from rate and spacecraft attitude sensors 241, 242 or other sensors within CMG 22, the ACP modifies the rate command issued to CMG 22 to control the acceleration and rate of the spacecraft. At the completion of the maneuver commands are issued to CMG 22 to return it to its quiescent state and stop the rotation of spacecraft 24. A similar mechanism permits exchange of momentum from the spacecraft to the CMG to counter external torque disturbances on the spacecraft. Under ordinary circumstances such an arrangement works well.
FIG. 2 is a simplified schematic block diagram of attitude control system 20 of FIG. 1 showing further details concerning CMG 22. Like reference numbers are used for like elements. CMG 22 comprises: (i) adder 30 which receives gimbal rate command (CRC) 46 from ACP 28 over link 29, (ii) error amplifier 32 having its input coupled to the output of adder 30 via link 31, (iii) filter 34 having its input coupled to the output of error amplifier 32 via link 33, (iv) limiter 36 having its input coupled to the output of filter 34 via link 35, (v) motor driver 38 whose input is coupled to the output of limiter 36 via link 37, and (vi) gimbal motor 40 whose input is coupled to the output of driver 38 via link 39. Filter 34 is used to compensate for vibrational resonance modes that may occur in the spacecraft so that the overall attitude control system is unconditionally stable. Limiter 36 insures that current drive Id to gimbal motor 40 does not exceed the maximum safe current Imax for driver 38 and/or motor 40. CMG 22 further comprises inner gimbal assembly 42 which is mechanically coupled to gimbal motor 40 and to spacecraft 24. Gimbal motor 40 delivers motor torque (MT) 41 to inner gimbal assembly 42. Inner gimbal assembly 42 delivers gyroscopic torque (GT) 25 to spacecraft 24. Inner gimbal assembly 42 can also receive reaction torque (RT) 27 from spacecraft 24 if spacecraft 24 is rotating. In this situation, the spacecraft is trying to turn inner gimbal assembly 42 instead of the other way around. tachometer 44 (abbreviated as “TACH”) is coupled to inner gimbal assembly 42 via link 43 and measures the rate of rotation of inner gimbal assembly 42 (and also indirectly the spacecraft rotation rate) and communicates measured rate feedback values (RFB) 49 back to ADDER 30 via feedback link 45. Elements 30, 32, 34, 36, 38 make up gimbal loop controller 23 which receives gimbal rate commands (GRC) 46 over link 29 from ACP 28 and gimbal rate feedback (RFB) 49 over link 45 from TACH 44 and delivers motor drive current Id 48 over link 39 to motor 40 so as to reduce the difference between gimbal rate command (GRC) 46 and rate feedback (RFB) 49 to zero, thereby supplying motor torque (MT) 41 to rotate the CMG gimbal at the desired rotation rate corresponding to GRC 46 and producing gyroscopic torque (GT) 25 acting on spacecraft 24. Gimbal Loop Controller 23 is desirably a Type 1 controller, that is, it includes summed integrating and linear amplifiers within error amplifier 32. This integrating action permits the control system to reject the effects of the reaction torque (RT) 27. This is conventional.
A problem with such systems is that if a spacecraft is moving as a result of prior gimbal rate command (GRC) 46 having been issued to CMG 22 from ACP 28, and ACP 28 or some other element of the system has a malfunction or reaches a time-out so that gimbal rate command (GRC) 46 being issued by ACP 28 drops to zero or otherwise becomes invalid, then spacecraft 24 and CMG 22 can be subjected to very severe mechanical stresses. For example, if GRC 46 is suddenly removed or becomes zero, then only RFB signal 49 is reaching adder 30 and gimbal loop control 23 reverses the direction of motor torque (MT) 41 as it tries to drive RFB 49 to zero. This can result in a sudden and potentially harmful deceleration of inner gimbal assembly 42 and spacecraft 24. Further, if driver 38 is disabled by command or as a failure mitigation response, the motor torque (MT) 41 will become zero but the reaction torque (RT) 27 remains, potentially resulting in a large acceleration of the gimbal in the opposite direction. This process can also result in damaging stresses to the CMG system and the spacecraft structure. Many high agility attitude control systems have historically used CMG's with geared drive motors, whose back-electro-motive-force (Vbemf) from motor 40 could be used in a dynamic braking circuit arrangement to passively limit the gimbal rate of a CMG caused by vehicle reaction torque. But some spacecraft employ non-geared, direct drive gimbal motors, where this approach is not practical. Hence, the problem is exacerbated with these vehicles. Accordingly, there continues to be a need for satellite active motion damping systems and methods that mitigate or avoid the physical stress, torque errors and other problems arising from various kinds of CMG gimbal drive interruptions.
Accordingly, it is desirable to provide improved active motion damping means and methods for satellites and other spacecraft. In addition, it is desirable that the apparatus and method be simple, rugged, reliable and require minimal change in the satellite hardware. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and the foregoing technical field and background.